High performance rocket engine having a stepped expansion combustion chamber and method of making the same

ABSTRACT

An improved rocket engine combustion chamber including a first chamber having a first diameter and located intermediate to a propellant injector and a second chamber having a second diameter that is larger than the first diameter. The combustion chamber extends radially outward from the first diameter to the second diameter suddenly at the intersection between the first and second chambers. Film cooling is provided by providing a stratified layer of low temperature fluid adjacent to the inner wall of the first chamber and surrounding a primary inner core of high temperature gases. The sudden stepped expansion at the interface between the first and second chambers provides secondary recirculation mixing of the propellants and facilitates complete combustion. In an additional aspect, the inner surface of the first chamber may be made of a material that has a high degree of thermal conductivity to minimize temperature gradients.

CROSS-REFERENCE TO RELATED APPLICATION

[0001] This application is a continuation of co-pending application Ser.No. 09/112,390, filed Jul. 9, 1998, the disclosure of which is herebyexpressly incorporated by reference and priority from the filing date ofwhich is hereby claimed under 35 U.S.C. § 120.

FIELD OF THE INVENTION

[0002] The present invention relates generally to the field of rocketengines and, more particularly, to an improved rocket engine combustionchamber design and method of making the same, wherein the combustionchamber has a first smaller diameter, film-cooled surface portionadjacent to a propellant injector and steps suddenly outward to a secondlarger-diameter portion at a position spaced away from the propellantinjector, wherein the film cooling, together with the sudden expansionof the diameter of the combustion chamber, results in an exceptionallyhigh degree of combustion efficiency.

BACKGROUND OF THE INVENTION

[0003] The field of rocket science has advanced rapidly during thelatter half of the Twentieth Century from its relatively primitivebeginnings. Early rockets were essentially experimental, pilotlessaircraft, which were operated by crude control systems. The tremendoustechnological advances in rocket propulsion have been accompanied bysimilar advances in other essential fields—such as electronics, inertialguidance and control systems, aerodynamics, and material sciences. As aresult, rockets today are manufactured for a variety of purposes,ranging from military applications to carrying scientific instrumentsfor use in gathering information at high altitudes, either within orabove the earth's atmosphere.

[0004] While such rockets may vary considerably, both in application aswell as size, they all include three essential components: a guidanceand control system, a mission payload that is to be carried by therocket, and a power source for propelling them. The first of thesecomponents is the guidance and control system that controls the flightpath of the rocket. The second of the aforementioned components is themission payload, which, as mentioned above, may vary widely, varyingfrom scientific instruments to surveillance equipment to explosivewarheads.

[0005] It is the third of the three essential components of arocket—namely, the power source—that is the focus of the presentinvention. This power source is typically a self-contained rocketengine. Three different types of rocket engines have been predominantlyutilized in the past—namely, solid propellant systems, liquidbipropellant systems, and liquid or gaseous monopropellant systems.Solid propellant systems present several significant disadvantages notfound in liquid bipropellant and monopropellant systems. For example,solid propellant systems are relatively heavy, have lower attainableexhaust velocities, and offer poor control of operating level in flight(throttleability).

[0006] Liquid bipropellant systems use an oxidizer and a fuel that aretanked separately and mixed in the combustion chamber. Typically, suchliquid bipropellant systems use hydrazine or monomethylhydrazine as thefuel and nitrogen tetroxide as the oxidizer. In some applications,bipropellant systems use gels instead of liquids. Liquid monopropellantsystems typically also use hydrazine as a monopropellant fuel. Sinceliquid bipropellant systems are more widely used, the discussion thatfollows focuses on such systems.

[0007] The typical components of a liquid bipropellant propulsion systemare the rocket engine, fuel tanks, and a vehicle structure to maintainthese parts in place and connect them to the mission payload. The liquidbipropellant rocket engine itself consists of a main chamber for mixingand burning the fuel and the oxidizer, with the fore end occupied byfuel and oxidizer manifolds and injectors, and the aft end comprising anozzle. The oxidizer and the fuel are transferred from their respectivetanks by pumps or may be pressurized by gas and are supplied to theinjector manifold at a high pressure. The oxidizer and the fuel are theninjected into the combustion chamber in a manner that assuresatomization and mixing so that they may be efficiently reacted toproduce thrust from the rocket engine.

[0008] Two problems that must be faced in the implementation of a rocketengine design are maximizing the efficiency of combustion and dealingwith the problem of heat in the rocket engine. It will at once beappreciated by those skilled in the art that it is desirable to have acombustion efficiency approaching as close as possible to 100%. Inaddition, those skilled in the art will also realize that the hot gaschemical and temperature environments of a rocket engine, if leftunchecked, may damage or destroy the combustion chamber during thedesired lifetime of the rocket engine (which is typically in the rangeof 35,000-100,000 seconds).

[0009] The main combustion chamber of larger rocket engines typicallyuses regenerative propellant cooling, in which the combustion chamberincludes a coolant jacket through which liquid propellant (usually fuel)is circulated at rates high enough to allow the rocket engine to operatecontinuously without an excessive increase in the combustion chamberwall temperature. Smaller rocket engines instead use direct rejection ofheat from the combustion chamber to the space environment by radiationheat transfer.

[0010] Effective cooling of a liquid rocket engine in the thrust rangeof 1 Newton to 10,000 Newtons is typically accomplished by using liquidor gaseous film cooling of the combustion chamber wall, whichestablishes a stratified layer of low temperature fluid adjacent to theinner wall of the combustion chamber. This is accomplished byestablishing a film cooling injection pattern and a main core injectionpattern, wherein the injectors provide a primary inner core of hightemperature gases and a peripheral layer of low temperature unmixed andpartially mixed propellant gases. The unmixed propellant used for thefilm cooling and partially mixed propellants must then be reacted in arapid and efficient manner in order to provide a maximum, specificimpulse efficiency rocket engine.

[0011] Several patents that are relevant to the present invention may bereviewed as background information. These patents are U.S. Pat. No.3,074,469, to Babbitt et al.; U.S. Pat. No. 4,785,748, to Sujata et al.;U.S. Pat. No. 4,915,038, also to Sujata et al.—all of which are assignedto the assignee of the present invention, as well as U.S. Pat. No.4,882,904, to Schoenman, and U.S. Pat. No. 4,936,091, also to Schoenman.U.S. Pat. Nos. 3,074,469; 4,785,748; 4,882,904; 4,915,038; and4,936,091, are each hereby incorporated herein by reference.

[0012] It is accordingly one of the principal objectives of the presentinvention that it result in a rocket engine having a design and methodof manufacture that provide a highly effective cooling mechanism, whichprotects the combustion chamber from damage or destruction caused byhigh temperature conditions. It is a further objective of the presentinvention that it minimizes or eliminates the reactions that take placebetween the incompletely reacted fuel and oxidizer products and thecombustion chamber wall materials. It is a related objective of thepresent invention that it optimizes the temperature gradients betweenthe various components of the rocket engine to provide effective coolingand minimize structural and thermal stresses.

[0013] It is another of the principal advantages of the presentinvention that it enhances the combustion efficiency of the rocketengine to the maximum degree possible. It is accordingly an objective ofthe present invention that the rocket engine combustion chamber be of adesign that promotes a complete mixing of the propellants such that theymay be completely reacted within the combustion chamber. It is a relatedobjective of the present invention that mixing of the main core of gaswith the film cooling layer is accomplished after the need for the filmcooling layer is no longer required, but before the unmixed andunreacted propellants leave the combustion chamber.

[0014] The stepped expansion combustion chamber rocket engine of thepresent invention must be of a construction that is both durable andlong lasting, and it must also require that no maintenance be providedby the user throughout its operating lifetime. In order to enhance themarket appeal of the stepped expansion combustion chamber rocket engineof the present invention, it should also be of relatively inexpensiveconstruction to thereby afford it the broadest possible market. Finally,it is also an objective that all of the aforesaid advantages andobjectives of the stepped expansion combustion chamber rocket engine ofthe present invention be achieved without incurring any substantialrelative disadvantage.

SUMMARY OF THE INVENTION

[0015] The disadvantages and limitations of the background art discussedabove are overcome by the present invention. With this invention, threekey aspects are incorporated into the design of a rocket enginecombustion chamber. The first two of these three key aspects are the useof film cooling and a stepped expansion combustion chamber, whichtogether provide the heretofore mutually exclusive benefits of effectivecooling of the combustion chamber and superior mixing of the fuel andoxidizer, resulting in highly efficient combustion.

[0016] The combustion chamber thus consists of two portions—namely, afirst portion referred to herein as a precombustion chamber, and asecond portion referred to herein as a main combustion chamber. Theprecombustion chamber has a first diameter and is located intermediateto the injector manifold assembly and the main combustion chamber, thelatter of which has a second diameter larger than the first diameter.The precombustion chamber and the main combustion chamber are coaxialand adjacent to each other, such that the combustion chamber extendsradially outward from the first diameter to the second diameter at theintersection between the precombustion chamber and the main combustionchamber in a step-wise manner. This construction is the derivation ofthe reference to a “stepped” combustion chamber.

[0017] The convergent throat and exhaust nozzle sections of the rocketengine form the remainder of the rocket engine. This section is formedin one single continuous assembly and is connected to the maincombustion chamber at the end opposite the precombustion chamber.

[0018] The injector manifold assembly contains fuel and oxidizermanifolds that are located therein, as well as injectors communicatingbetween the respective fuel and oxidizer manifolds and the interior ofthe stepped combustion chamber. The fuel manifold will be supplied withpressurized fuel, while the oxidizer manifold will be supplied withpressurized oxidizer. The injectors establish two spray patterns intothe stepped combustion chamber, namely, a main core injection patternand a film-cooling pattern. The main core injection pattern will providea primary inner core of well mixed, high temperature, combusting gases,while the film-cooling pattern will provide an annular peripheral layerof low temperature unmixed and partially mixed propellant gasesimmediately adjacent to the interior surface of the precombustionchamber.

[0019] It will therefore be appreciated by those skilled in the art thatonly the injector manifold assembly and the precombustion chamber comein contact with the oxidizer and fuel and partially reacted combustionproducts at low temperatures. The precombustion chamber is effectivelycooled by film cooling, and the sudden expansion process effectivelymixes the remaining fuel and oxidizer, allowing them to combustcompletely. The injector manifold assembly and the precombustion chambercome in contact with the oxidizer and fuel, partially reacted combustionproducts and, in some cases, decomposing fuel—but only at relatively lowtemperatures. The larger diameter main combustion chamber is subjectedto higher temperatures, but only in the presence of the fully combustedpropellants. Thus, the entire stepped expansion combustion chamber ofthe rocket engine of the present invention is protected from beingsimultaneously exposed to both corrosive, partially reacted combustionproducts and high temperatures.

[0020] The use of the sudden expansion design in the stepped expansioncombustion chamber of the rocket engine of the present invention alsoenhances the combustion efficiency by promoting mixing of the main coreof gas with the film layer after the need for the cooling effectprovided by the film layer is no longer required. The sudden expansiondesign is also effective in providing a flame-holding and recirculationzone to increase the chamber residence time of the unreacted andpartially reacted gases. The resultant momentum of the main core gasesis designed to impinge on the precombustion chamber wall just prior tothe sudden change in diameter of the combustion chamber. The mixed coreand film are thus effectively reacted in the recirculation zone, whichresults from the sudden dimensional expansion.

[0021] The third key aspect of the stepped expansion combustion chamberrocket engine of the present invention is the use of a material for theinner surface of the precombustion chamber that has a high degree ofthermal conductivity. The use of a material having a relatively highthermal conductivity for the precombustion chamber will serve tominimize the wall axial and circumferential temperature gradient betweenthe portion of the combustion chamber at which the sudden expansionoccurs and the face of the injector manifold assembly. The use ofmoderate amounts of propellant as a film coolant effectively cools theinjector manifold assembly and eliminates the need for thermal isolationof the injector manifold assembly and the combustion chamber.

[0022] The use of a material for the inner surface of the precombustionchamber that has a high degree of thermal conductivity may befacilitated in two ways. First, the precombustion chamber itself may bemade of a material having a high degree of thermal conductivity.Alternately, an inner precombustion chamber liner made of a materialhaving a high degree of thermal conductivity may be used. The latterapproach has an advantage in that the inner precombustion chamber linercan be fitted inside the precombustion chamber, with the precombustionchamber itself being made of a material that is better suited forassembly, together with the injector manifold assembly and the maincombustion chamber.

[0023] It may therefore be seen that the present invention creates arocket engine having a design and method of manufacture that provide ahighly effective cooling mechanism, which protects the combustionchamber from damage or destruction caused by high temperatureconditions. The present invention minimizes or eliminates the reactionsthat take place between the incompletely reacted fuel and oxidizerproducts and the combustion chamber wall materials. In a related aspect,the stepped expansion combustion chamber rocket engine of the presentinvention optimizes the temperature gradients between the variouscomponents of the rocket engine, thereby providing effective cooling andminimizing structural and thermal stresses.

[0024] The stepped expansion combustion chamber rocket engine of thepresent invention also enhances the combustion efficiency of the rocketengine to the maximum degree possible. The stepped expansion combustionchamber is of a design that promotes a complete mixing of thepropellants such that they may be completely reacted within thecombustion chamber. In the stepped expansion combustion chamber rocketengine of the present invention, mixing of the main core of gas with thefilm cooling layer is accomplished after the need for the film coolinglayer is no longer required but before the unmixed and unreactedpropellants leave the combustion chamber.

[0025] The stepped expansion combustion chamber rocket engine of thepresent invention is of a construction that is both durable and longlasting, and which will require no maintenance to be provided by theuser throughout its operating lifetime. The stepped expansion combustionchamber rocket engine of the present invention is also of relativelyinexpensive construction to enhance its market appeal and to therebyafford it the broadest possible market. Finally, all of the aforesaidadvantages and objectives of the stepped expansion combustion chamberrocket engine of the present invention are achieved without incurringany substantial relative disadvantage.

BRIEF DESCRIPTION OF THE DRAWINGS

[0026] The foregoing aspects and many of the attendant advantages ofthis invention will become more readily appreciated as the same becomebetter understood by reference to the following detailed description,when taken in conjunction with the accompanying drawings, wherein:

[0027]FIG. 1 is a somewhat schematic, cross-sectional view from the sideof a stepped expansion combustion chamber constructed according to theteachings of the present invention, showing a plurality of injectorslocated in an injector manifold assembly, a smaller diameterprecombustion chamber, and a larger diameter main combustion chamber;

[0028]FIG. 2 is a cross-sectional view from the side of a portion of astepped expansion combustion chamber, also constructed according to theteachings of the present invention, showing an inner precombustionchamber liner made of a material having a high thermal conductivitybeing located within an outer precombustion chamber sleeve; and

[0029]FIG. 3 is a partial cross-sectional view from the side of a rocketengine using the stepped expansion combustion chamber rocket engine ofthe present invention, which is illustrated in FIG. 1.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

[0030] The preferred embodiment of the stepped expansion combustionchamber rocket engine of the present invention is illustrated in FIG. 1,which shows the four principal components that together define thecombustion chamber. A hollow, essentially cylindrical precombustionchamber 20 having a first diameter is mounted onto a hollow, essentiallycylindrical main combustion chamber 22 having a second diameter that islarger than the first diameter, using a washer-shaped annular step 24,which is located intermediate the precombustion chamber 20 and the maincombustion chamber 22.

[0031] The inner diameter of the annular step 24 is the same as theinner diameter of the precombustion chamber 20, and the outer diameterof the annular step 24 is the same as the outer diameter of the maincombustion chamber 22. The precombustion chamber 20 and the maincombustion chamber 22 are thereby mounted into a single unit in coaxialfashion, preferably by welding, but by any method that ensures aleak-tight connection, such as brazing, swaging, and explosive forming.The end of the precombustion chamber 20 extending away from the maincombustion chamber 22 has a flange 26, which extends radially outwardlytherefrom.

[0032] Mounted onto the flange 26 on the end of the precombustionchamber 20 extending away from the main combustion chamber 22 is aninjector manifold assembly 30. The injector manifold assembly 30 ispreferably mounted onto the precombustion chamber 20 by welding. Adisk-shaped first manifold 32 is centrally located within the injectormanifold assembly 30. An annular second manifold 34 is also locatedwithin the injector manifold assembly 30, with the second manifold 34being spaced outwardly away from the first manifold 32. The firstmanifold 32 and the second manifold 34 are coaxial with each other andwith the precombustion chamber 20 and the main combustion chamber 22.

[0033] A first propellant supply channel 36 will be used to supply afirst propellant to the first manifold 32. Similarly, a secondpropellant supply channel 38 will be used to supply a second propellantto the second manifold 34. The first propellant supply channel 36 andthe second propellant supply channel 38 will communicate with apparatusto supply first and second propellants, respectively, to the firstmanifold 32 and the second manifold 34, respectively. The firstpropellant supply channel 36 and the second propellant supply channel 38thus communicate with the exterior of the injector manifold assembly 30on the side of the injector manifold assembly 30 opposite theprecombustion chamber 20 and the main combustion chamber 22.

[0034] A plurality of injector channels communicate between the firstmanifold 32 and the second manifold 34 and the side of the injectormanifold assembly 30 facing inwardly into the precombustion chamber 20and the main combustion chamber 22. Specifically, a plurality of firstinjector channels 40 extends in an outwardly conical frustrum-shapedarray from the first manifold 32 to the side of the injector manifoldassembly 30 facing inwardly into the precombustion chamber 20 and themain combustion chamber 22. A plurality of second injector channels 42extends in an inwardly conical frustrum-shaped array from the firstmanifold 32 to the side of the injector manifold assembly 30 facinginwardly into the precombustion chamber 20 and the main combustionchamber 22.

[0035] The points at which the first injector channels 40 and the secondinjector channels 42 open into the precombustion chamber 20 form aroughly circular array of openings at approximately the midpoint ofradii of the face of the injector manifold assembly 30 facing inwardlyinto the precombustion chamber 20 and the main combustion chamber 22. Inthis circular array, openings from the first injector channels 40 andthe second injector channels 42 alternate to form a main core injectorpattern. Multiple spray angles may be used if so desired. First andsecond propellants injected from the first injector channels 40 and thesecond injector channels 42, respectively, will mix well, and provide aprimary inner core of high temperature combusting gases.

[0036] A plurality of third injector channels 44 extends in an outwardlyconical frustrum-shaped array from the second manifold 34 to the side ofthe injector manifold assembly 30 facing inwardly into the precombustionchamber 20 and the main combustion chamber 22. The points at which thethird injector channels 44 open into the precombustion chamber 20 form aroughly circular array of openings close to the intersection of theinner surface of the precombustion chamber 20 and the surface of theinjector manifold assembly 30, which faces inwardly into theprecombustion chamber 20 and the main combustion chamber 22. In thiscircular array, openings from the third injector channels 44 form aperipheral layer injection pattern.

[0037] This peripheral layer injection pattern of low temperatureunmixed second propellant comprises the film cooling flow, which willact to cool the inner surface of the precombustion chamber 20. Thiscylindrical, film cooling layer will thus extend from the openings fromthe third injector channels 44 along the length of the precombustionchamber 20. Typically, the first propellant is the oxidizer and thesecond propellant, which is used for the film cooling flow, is the fuel.Thus, while the inner surface of the precombustion chamber 20 does comeinto contact to some extent with the oxidizer and fuel and withpartially reacted combustion products, only relatively low temperaturesexist at this inner surface of the precombustion chamber 20.

[0038] The sudden stepped expansion at the interface between theprecombustion chamber 20 and the main combustion chamber 22 will providesecondary recirculation mixing of the propellants and will completecombustion between the main hot gas core created by the combustion ofthe propellants coming from the first injector channels 40 and thesecond injector channels 42 on the one hand, and the propellant filmcooling layer coming from the third injector channels 44, on the otherhand. This complete combustion occurs near the entrance to the maincombustion chamber 22, which is subjected to higher temperatures, butonly in the presence of the fully combusted propellants.

[0039] Thus both the precombustion chamber 20 and the main combustionchamber 22 are protected from being simultaneously exposed to bothcorrosive partially reacted combustion products and high temperatures.The stepped expansion combustion chamber of the rocket engine of thepresent invention greatly enhances the combustion efficiency bypromoting mixing of the main core of gas with the film layer after theneed for the cooling effect provided by the film layer is no longerrequired. Combustion efficiencies of over 99.6% have been obtained usingthe stepped expansion combustion chamber rocket engine of the presentinvention.

[0040] Use of this invention is applicable to, but not limited to, thefollowing liquid and gaseous propellants—nitrogen tetroxide (N₂O₄) andmixtures of N₂O₄ containing nitric oxide (NO), red fuming nitric acidincluding all inhibitors, oxygen (O₂), ozone, fluorine and chlorine andcompounds and mixtures thereof, hydrazine and all related compounds thatinclude carbon and hydrogen radicals, including monomethylhydrazine, airand mixtures thereof containing oxygen, unsymmetrical dimethylhydrazineand mixtures thereof, hydrogen, methane, ethanol, propane and relatedcarbon/hydrogen and carbon/hydrogen/oxygen compounds.

[0041] Multiple materials may be used for the precombustion chamber 20,the main combustion chamber 22, the annular step 24, and the injectormanifold assembly 30 to minimize stress loads resulting from differencesin the thermal coefficients of expansion of materials. In the preferredembodiment, the precombustion chamber 20 is made of a material that hasa high degree of thermal conductivity to minimize localized, abnormallyhigh heating rates and temperature gradients. Examples of such materialsare the elements and alloys of nickel, platinum, rhodium, iridium,rhenium, beryllium, beryllium copper, and columbium (niobium). In thepreferred embodiment, the main combustion chamber 22 is made of amaterial that is capable of withstanding high temperatures. Examples ofsuch materials are the elements and alloys of rhenium, iridium platinum,rhodium, carbon, and silicon carbide.

[0042] One problem that may occur in using a material that has a highdegree of thermal conductivity is that it may be difficult or impossibleto use a material for the precombustion chamber 20 that cannot be weldedto the injector manifold assembly 30 or to the main combustion chamber22 via the annular step 24 because of metallurgical incompatibility. Itis possible, however, to use a material that has a high degree ofthermal conductivity by cladding, diffusion bonding, brazing, alloying,or interference fitting it to another material that is capable of beingwelded to the injector manifold assembly 30 or to the main combustionchamber 22 via the annular step 24. Alternately, an interface materialthat provides minimum thermal contact resistance may be placedintermediate to the outer precombustion chamber sleeve 50 and the innerprecombustion chamber liner 52, such that the interface material isre-melted or forced by a pressure load to displace any void between theouter precombustion chamber sleeve 50 and the inner precombustionchamber liner 52.

[0043] Referring next to FIG. 2, such a use of material that has a highdegree of thermal conductivity is illustrated. A hollow, essentiallycylindrical outer precombustion chamber sleeve 50 is illustrated, whichhas a hollow, essentially cylindrical inner precombustion chamber liner52 fixedly mounted therein. The inner precombustion chamber liner 52 isformed to be inserted into the outer precombustion chamber sleeve 50.Optionally, the outer diameter of the inner precombustion chamber liner52 and the inner diameter of the outer precombustion chamber sleeve 50may be tapered to fit together, as is shown in FIG. 2. The outerprecombustion chamber sleeve 50 and the inner precombustion chamberliner 52 are assembled together using any of the aforementionedtechniques.

[0044] A main combustion chamber 54 is mounted onto one end of the outerprecombustion chamber sleeve 50 by a welded joint 56 therebetween. Theinner diameter of the inner precombustion chamber liner 52 is smallerthan the inner diameter of the main combustion chamber 54. The edge ofthe inner precombustion chamber liner 52 extends somewhat into the maincombustion chamber 54 and forms a stepped expansion therein. An injectormanifold assembly 58 is mounted onto the other end of the outerprecombustion chamber sleeve 50 by a welded joint 60 therebetween.

[0045] The inner precombustion chamber liner 52 is made of a materialthat has a high degree of thermal conductivity, but which cannotnecessarily be welded to the main combustion chamber 54 or the injectormanifold assembly 58. Examples of preferred materials for the innerprecombustion chamber liner 52 include platinum, platinum alloys withrhodium, iridium alloys with rhodium, niobium, nickel alloys, cobaltalloys, carbon, carbon composites, and silicon carbide—all of which arecapable of withstanding an operating temperature in excess of 2000° F.

[0046] Referring finally to FIG. 3, the combustion chamber illustratedin FIG. 1, which is identified by the reference numeral 70, is shownassembled together with several additional components to form thestepped expansion combustion chamber rocket engine of the presentinvention. A convergent section, throat, and nozzle 72, are shown asbeing mounted on the open end of the main combustion chamber 22 of thecombustion chamber 70. A first propellant valve 74 controls the flow ofa first propellant to the combustion chamber 70, and a second propellantvalve 76 controls the flow of a second propellant to the combustionchamber 70. Completing the construction is a mounting bracket 78, whichmay be used to mount the stepped expansion combustion chamber rocketengine of the present invention onto a rocket (not shown).

[0047] The capabilities and performance of the stepped expansioncombustion chamber rocket engine of the present invention are quiteimpressive. The stepped expansion combustion chamber will operate at alevel that is below the melting point of the material and at a materialoxidation level that is consistent with operational times in the rangeof 35,000-40,000 seconds. The rocket engine will operate at any desiredon and off times without overheating the rocket engine itself, theinjectors, or the propellant valves. The stepped expansion combustionchamber rocket engine of the present invention is able to meet thespecific stress margins required to prevent failure of the combustionchamber due to differences in the coefficients of thermal expansion ofthe materials and minimize radial and circumferential temperaturegradients by using high thermal conductivity materials. All of this isattained while also achieving near theoretical specific impulseperformance levels.

[0048] It may therefore be appreciated from the above detaileddescription of the preferred embodiment of the present invention that itcreates a rocket engine having a design and method of manufacture thatprovide a highly effective cooling mechanism that protects thecombustion chamber from damage or destruction caused by high temperatureconditions. The present invention minimizes or eliminates the reactionsthat take place between the incompletely reacted fuel and oxidizerproducts and the combustion chamber wall materials. In a related aspect,the stepped expansion combustion chamber rocket engine of the presentinvention optimizes the temperature gradients between the variouscomponents of the rocket engine, thereby providing effective cooling andminimizing structural and thermal stresses.

[0049] The stepped expansion combustion chamber rocket engine of thepresent invention also enhances the combustion efficiency of the rocketengine to the maximum degree possible. The stepped expansion combustionchamber is of a design that promotes a complete mixing of thepropellants, such that they may be completely reacted within thecombustion chamber. In the stepped expansion combustion chamber rocketengine of the present invention, mixing of the main core of gas with thefilm cooling layer is accomplished after the need for the film coolinglayer is no longer required, but before the unmixed and unreactedpropellants leave the combustion chamber.

[0050] The stepped expansion combustion chamber rocket engine of thepresent invention is of a construction that is both durable and longlasting, and which will require no maintenance to be provided by theuser throughout its operating lifetime. The stepped expansion combustionchamber rocket engine of the present invention is also of relativelyinexpensive construction to enhance its market appeal and to therebyafford it the broadest possible market. Finally, all of the aforesaidadvantages and objectives of the stepped expansion combustion chamberrocket engine of the present invention are achieved without incurringany substantial relative disadvantage.

[0051] Although an exemplary embodiment of the stepped expansioncombustion chamber rocket engine of the present invention has been shownand described with reference to particular embodiments and applicationsthereof, it will be apparent to those having ordinary skill in the artthat a number of changes, modifications, or alterations to the inventionas described herein may be made, none of which depart from the spirit orscope of the present invention. All such changes, modifications, andalterations should therefore be seen as being within the scope of thepresent invention.

[0052] While the preferred embodiment of the invention has beenillustrated and described, it will be appreciated that various changescan be made therein without departing from the spirit and scope of theinvention.

The embodiments of the invention in which an exclusive property orprivilege is claimed are defined as follows:
 1. A rocket enginecomprising: a precombustion chamber having a first end and a second end,said precombustion chamber being defined by a peripheral wall extendingbetween said first and second ends of said precombustion chamber andhaving a first inner diameter along a majority of said peripheral wall;a main combustion chamber having a first end and a second end, said maincombustion chamber being defined by a peripheral wall extending betweensaid first and second ends of said main combustion chamber and having asecond inner diameter which is larger than said first inner diameter;and an annular step interconnecting said second end of saidprecombustion chamber and said first end of said main combustion chamberinto a combustion chamber assembly, said first end of said precombustionchamber and said second end of said main combustion chamber extending inopposite directions, the inner diameter of said combustion chamberassembly stepping suddenly outward from said first inner diameter insaid precombustion chamber to said second inner diameter in said maincombustion chamber at the location of said annular step.
 2. A rocketengine as defined in claim 1, wherein said precombustion chamber is of ahollow, essentially cylindrical configuration having said first innerdiameter.
 3. A rocket engine as defined in claim 1, wherein at least aninner surface of said precombustion chamber is made of a material havinga high degree of thermal conductivity.
 4. A rocket engine as defined inclaim 1, wherein said precombustion chamber is made of a material fromthe group consisting of the elements and alloys of nickel, platinum,rhodium, iridium, rhenium, beryllium, copper, and columbium.
 5. A rocketengine as defined in claim 1, wherein said precombustion chambercomprises: a sleeve having a first end and a second end, at least oneend being interconnected to said first end of said main combustionchamber; and a liner located inside said sleeve, an inside surface ofsaid liner defining said first inner diameter of said precombustionchamber.
 6. A rocket engine as defined in claim 5, wherein said liner ismade of a material having a high degree of thermal conductivity.
 7. Arocket engine as defined in claim 6, wherein said liner is made of amaterial from a group consisting of platinum, platinum alloys withrhodium, iridium, iridium alloys with rhodium, niobium, nickel alloys,cobalt alloys, carbon, carbon composites, and silicone carbide.
 8. Arocket engine as defined in claim 5, wherein a portion of said linerextends partially into said main combustion chamber, said portion ofsaid liner that extends into said main combustion chamber having an endsurface that comprises the annular step.
 9. A rocket engine as definedin claim 5, wherein said second end of said sleeve, said annularsurface, and said first end of said main combustion chamber are joinedtogether by one of the techniques from a group consisting of welding andbrazing, and wherein said first end of said outer precombustion chambersleeve is joined to said injector manifold assembly by one of thetechniques from a group consisting of welding and brazing.
 10. A rocketengine as defined in claim 5, wherein said liner is installed insidesaid sleeve by one of the techniques from a group consisting ofcladding, diffusion bonding, brazing, alloying, interference fitting,swaging and explosive forming and bonding.
 11. A rocket engine asdefined in claim 5, wherein said liner is installed inside said outerprecombustion chamber sleeve by placing an interface material thatprovides minimum thermal contact resistance intermediate said innerprecombustion chamber liner and said outer precombustion chamber sleevesuch that said interface material is re-melted or forced by a pressureload to displace any void between said inner precombustion chamber linerand said outer precombustion chamber sleeve.
 12. A rocket engine asdefined in claim 5, wherein an outer surface of said liner and an innersurface of said sleeve are each correspondingly tapered so that theouter surface of said liner substantially engages the inner surface ofsaid sleeve when the liner is located within said sleeve.
 13. A rocketengine as defined in claim 1, wherein said main combustion chamber is ofa hollow, essentially cylindrical configuration having said second innerdiameter.
 14. A rocket engine as defined in claim 1, wherein said maincombustion chamber is made of a material from a group consisting of theelements and alloys of rhenium, iridium, platinum, rhodium, carbon, andsilicone carbide.
 15. A rocket engine as defined in claim 1, whereinsaid second end of said precombustion chamber is installed inside saidfirst end of said main combustion chamber, thereby forming said annularstep.
 16. A rocket engine as defined in claim 15, wherein said secondend of said precombustion chamber is connected to said first end of saidmain combustion chamber by one of the techniques from a group consistingof welding and brazing.
 17. A rocket engine as defined in claim 1,wherein said inner diameter of said combustion chamber is substantiallyconstant from said first end of said precombustion chamber to saidsecond end of said precombustion chamber.
 18. A rocket engine combustionchamber comprising: a first chamber portion defining a first innerdiameter for a majority of the inner portion length; a second chamberportion defining a second inner diameter, wherein the second innerdiameter is greater than the first inner diameter; and an annularportion extending from the first inner diameter to the second innerdiameter.
 19. The rocket engine combustion chamber of claim 18, whereinthe inner surface of the first chamber portion defines a hollow,essentially cylindrical cavity having the first inner diameter.
 20. Therocket engine combustion chamber of claim 18, wherein at least the innersurface of the first chamber portion is made of a material having a highdegree of thermal conductivity.
 21. The rocket engine combustion chamberof claim 18, wherein the first chamber portion is made of a materialfrom a group consisting of the elements and alloys of nickel, platinum,rhodium, iridium, rhenium, beryllium, copper, and columbium.
 22. Therocket engine combustion chamber of claim 18, wherein the first chamberportion further comprises: a sleeve having a first end and a second end,at least one end being interconnected to the second chamber portion; anda liner disposed inside of the sleeve, the liner having an insidesurface defining the first inner diameter of the first chamber portion.23. The rocket engine combustion chamber of claim 22, wherein the linercomprises a material having a high degree of thermal conductivity. 24.The rocket engine combustion chamber of claim 23, wherein the liner ismade of a material from a group consisting of platinum, platinum alloyswith rhodium, iridium, iridium alloys with rhodium, niobium, nickelalloys, cobalt alloys, carbon, carbon composites, and silicone carbide.25. The rocket engine combustion chamber of claim 22, wherein a portionof the liner extends partially into the second chamber portion, whereinthe portion of the liner which extends into the second chamber portionhas an end surface that comprises the annular portion.
 26. The rocketengine combustion chamber of claim 22, wherein the second end of thesleeve, the annular portion, and the second chamber portion are joinedtogether by one of the techniques from a group consisting of welding andbrazing.
 27. The rocket engine combustion chamber of claim 22, whereinthe liner is installed inside the sleeve by one of the techniques from agroup consisting of cladding, diffusion bonding, brazing, alloying,interference fitting, swaging and explosive forming and bonding.
 28. Therocket engine combustion chamber of claim 22, wherein the liner isinstalled inside the sleeve by placing an interface material thatprovides minimum thermal contact resistance intermediate the liner andthe sleeve such that the interface material is re-melted or forced by apressure load to displace any void between the liner and the sleeve. 29.The rocket engine combustion chamber of claim 22, wherein an outersurface of the liner and an inner surface of the sleeve are eachcorrespondingly tapered so that the outer surface of the liner and theinner surface of the sleeve correspondingly engage each other when theliner is installed inside the sleeve.
 30. The rocket engine combustionchamber of claim 18, wherein the second chamber portion is of a hollow,essentially cylindrical configuration having the second inner diameter.31. The rocket engine combustion chamber of claim 18, wherein the secondchamber portion is made of a material from a group consisting of theelements and alloys of rhenium, iridium, platinum, rhodium, carbon, andsilicone carbide.
 32. The rocket engine combustion chamber of claim 18,wherein the first chamber portion is partially installed within thesecond chamber portion to thereby form the annular portion.
 33. Therocket engine combustion chamber of claim 32, wherein the first chamberportion is connected to the second chamber portion by one of thetechniques from a group consisting of welding and brazing.
 34. A rocketengine comprising: a first combustion chamber portion defining a firstinner diameter for a majority of the first combustion chamber portion; asecond combustion chamber portion defining a second inner diameter,wherein the second diameter is greater than the first diameter; and anannular portion extending from the first diameter to the seconddiameter.
 35. A rocket engine comprising: a first combustion chamberportion defining a first inner diameter for a majority of the firstcombustion chamber portion; a second combustion chamber portion defininga second inner diameter, wherein the second inner diameter is greaterthan the first inner diameter; an annular portion extending from thefirst inner diameter to the second inner diameter; and an injectionassembly coupled to the first combustion chamber portion.
 36. A rocketengine comprising: a first combustion chamber portion defining a firstinner diameter for a majority of the first combustion chamber portion; asecond combustion chamber portion defining a second inner diameter,wherein the second inner diameter is greater than the first innerdiameter; an annular portion extending from the first inner diameter tothe second inner diameter; an injection assembly coupled to the firstcombustion chamber portion; a throat coupled to the second combustionchamber portion; and a nozzle coupled to the throat.
 37. The rocketengine of claim 36, having a combustion efficiency of greater than orequal to 99.6%.
 38. A method of making a rocket engine combustionchamber comprising forming a first inner diameter for a majority of afirst inner chamber portion length, and forming an annular portionextending from the first diameter to a second chamber portion having asecond inner diameter, wherein the second inner diameter is greater thanthe first inner diameter.
 39. A method of making a rocket enginecomprising: forming a combustion chamber by coupling a first chamberdefined by a substantially constant first inner diameter to a secondchamber defined by a substantially constant second and greater innerdiameter in an end to end relationship, thereby forming an outwardlyexpansive annular portion at a location where the first chamber iscoupled to the second chamber.